1. Field of the Invention
The present invention relates, generally, to gas turbine engines. More particularly, the invention relates to the impeller of an axi-centrifugal compressor for gas turbine engines for aircraft. The invention has particular utility for improving the efficiency of a gas turbine engine by allowing a higher compressor discharge temperature.
2. Background Information
Though it does not depict any existing engine, FIG. 1 illustrates the current state-of-the-art for axi-centrifugal gas turbine aircraft engines, and is included to provide a frame of reference for the subsequent discussion of prior art and for the present invention. The direction is to the left in the figure, and aft is to the right. Axi-centrifugal gas turbine engines are very compact and efficient. The main airflow goes through a series of axial compressor stages 10 then through the impeller 12 which has a plurality of blades 14 which redirect the flow radially with centrifugal force into diffuser pipes 16 which increase the pressure and reduce the velocity of the airflow as it is redirected toward combustors 18. In the combustors, the air is mixed with fuel, ignited, and the resulting gas passed through blades of high-pressure turbines 20. A small portion of the main airflow, called cooling air bleed, is removed from the main airflow in front of impeller 12 and is directed afterward along the hub of the impeller to the high-pressure turbines 20 where it used to cool the blades 21 of the second stage high-pressure turbine before reentering the main airflow stream.
Referring also to FIG. 2, the impeller 12 is almost always made of titanium rather than steel due to titanium's higher strength to density ratio, which makes it ideal for rotating machinery components. Furthermore, titanium is much less expensive to purchase and machine than high strength steels. However, at sustained temperatures above 1000.degree. F., the strength of titanium diminishes rapidly with increasing temperature. With current titanium impellers, such as is illustrated in FIG. 2, the maximum compressor discharge temperature, usually identified by the symbol T3, is limited to 1100.degree. F.
The impeller temperatures are non-uniform. The peak temperature occurs near the rim 24 on the back face 22 at point 26 where radiant heat from the turbines 20 is reflected forward. Also, there is leakage of the hot main airflow around rim 24 onto the back face 22 of impeller 12, further exacerbating the heating of back face 22. The temperature at point 26 is approximately 150.degree. F. higher than anywhere else on impeller 12 at the high-power engine conditions. If, at those conditions, the temperature of the impeller at point 26 could be reduced 150.degree. F., that would allow the compressor discharge temperature T3 to be increased by 150.degree. F. to 1250.degree. F., thereby significantly increasing the overall engine efficiency.
The temperature at point 26 cannot be reduced simply by blowing cooling air at the back face 22 near rim 24 because the main airflow crosses a gap between the rim 24 and the diffuser pipes 16. The flow parameters across this gap are critical. Any cooling air directed at the back face 22 near rim 24 of impeller 12 would impinge on the main airflow and disrupt that critical flow sufficiently to destroy the effectiveness of the airflow into and through the diffuser pipes 16, resulting in a drastic reduction of engine efficiency.
U.S. Pat. No. 4,793,772 to Zaehring and U.S. Pat. Nos. 4,920,741 and 4,961,309, both to Liebl, disclose circulating cooling air in a chamber formed outside of the stub shaft to cool the last compressor section of an axial compressor. U.S. Pat. No. 4,808,073 to Zaehring et al. discloses vane-like ribs on the inside of the rear stub shaft which direct cooling air from the center shaft outwardly along the stub shaft and against the outer portion of the last rotor disk. These devices and methods work because the stub shaft connects to the last stage compressor rotor near the rim of the rotor. The shaft that connects to an impeller of an axi-centrifugal compressor connects near its hub rather than its rim, therefore, these devices and methods are not applicable to an impeller for an axi-centrifugal compressor.
The present invention provides an improved impeller for an axi-centrifugal gas turbine and a method of cooling it which reduces the temperature near the outer rim 150.degree. F. over conventional impellers without disrupting the critical airflow between the impeller and the diffuser pipes.